Miniature rocket



July 21, 1970 B. B, GouLD ETAL. 3,521,564

MINIATURE ROCKET Original Filed Feb. 11. 1965 2 Sheets-Sheet 1 "FIG/JE.

5:2 FIG. 5 54 FIG-.11.

- INVENToRs WILLLAM D. BARTON BERT .5. Gouna ARTHUR 11:51am,

BY RQBERT AHIHARDT www July 21, 1970 B. B. GouLD ETAL 3,521,554

' MINIATURE ROCKET Original Filed Feb. 11, 1965 2 Sheets-Sheet 3 Irma.. v

INVENTORS ROBERT vLNHARDT M "ZAgNEY United States Patent O 3,521,564 MINIATURE ROCKET Bert B. Gould, Alameda County, and Arthur T. Biehl and Robert Mainhardt, Contra Costa County, Calif., and William D. Barton, Panama, Republic of Panama, assignors to MB Associates, a corporation of California Original application Feb. 11, 1965, Ser. No. 435,780. Divided and this application Oct. 22, 1965, Ser. No. 515,513

Int. Cl. F42b 13/28 U.S. Cl. 10249.7 7 Claims This invention relates to miniature rockets suitable for anti-personnel use and other uses and it relates more particularly to a low cost, light weight self-propelled rocket of small dimension capable of attaining such high velocity as to enane 1t to be used as an extremely efficient kinetic energy kill mechanism or for use as other weaponry.

This application is a divisional application of my copending application Ser. No. 435,780, filed Feb. 11, 1965, entitled Small Arms Weapon, which is a continuationin-part of application Ser. No. 141,237, liled Sept. 20, 1961, which was a continuation-in-part of application Ser. No. 61,017, filed Oct. 6, 1960, the latter two applications now abandoned.

This invention relates to a novel ordnance weapon and methods for constructing the same. According to the basic concept of the weapon, a small rocket is provided for use as an antipersonnel kinetic energy kill mechanism. In accordance with various embodiments of the invention a plurality of such miniature rockets may be packaged together in a firing tube so that many may ybe fired simultaneously, or a number of such rockets may be dispersed and fired from a carrier such as a shell, bomb, or missile.

Traditionally, the rifle has been used as the standard small arms weapon. However, it suffers from a number of disadvantages. It is expensive to manufacture. The accuracy required in machining the barrel, and the precision required in the parts for feeding and aligning the bullet and cartridge, are costly such that providing a means for tiring a bullet is considered to be relatively expensive. A rifle is heavy both in itself and with respect to its ammunition, and it requires considerable skill to manipulate with any degree of effectiveness. The length of the barrel and the precision and rifling embodied in the barrel lining obviously require the riile to be of a size which makes concealment difficult, and calls for a rie of substantial weight, thereby imposing burdens on personnel using it. Further, the bullet has the undesirable characteristic of having its maximum velocity at the muzzle of the gun with continual deceleration thereafter, whereas one would prefer to have the maximum velocity in the probable target zone. Since it is necessary that a bullet must acquire its velocity from the expansion of gases while in the barrel of the rie, and prior to ejection from the gun, it is essential for the proper design of a rie to provide a barrel of sufficient length to enable full acceleration of the bullet before it emerges from the barrel. The barrel must be formed with suitable rilling to direct the bullet accurately at the time it is ejected from the barrel toward the target. The velocity of the rie bullet from the very start is too low for maximum effectiveness since, using ordinary techniques, the velocity is about 2,800 feet per second. If velocities of 3,00() feet per second and higher can be achieved, the victim suffers severe shock and tissue damage as contrasted to mere penetration achieved by an ordinary rie. Even less effectiveness is experienced in proportion to the distance the bullet travels from the rie before striking a target. Although attempts have been made to obtain higher muzzle velocities for ries, such attempts have invariably resulted in increased weapon weight and rapid erosion of the gun barrel.

3,521,564 Patented July 21, 1970 ICC Further, the rate of acceleration required to be effected within the length of the rie barrel introduces a considerable amount of recoil in the rifle. The user is thereby exposed to a rather large kick-back which makes it difficult to control the barrel and its effective use is limited to persons capable of absorbing the shock. Further, because of the high rate of acceleration and rotation of the bullet projected from the rie, it is diilicult to bond or join other elements with the bullet, such as CBR (chemical, biological. or radiological) agents, which also must be easily disseminated in the target, for delivery with the bullet to a distant target. T his factor therefore limits the rie to uses for delivery of bullets and like payloads.

Still further, the riile is subject to the deficiency that once the bullet is free of the barrel, it thereafter loses velocity such that maximum effectiveness, from the standpoint of speed or velocity, occurs closest to the rifle, whereas at distances 200 to 300 yards from the rifle, the velocity of the bullet has decreased to such extent as to no longer have ultimate effectiveness.

In contrast to the rifle and bullet, self-propelled miniature rockets, to which this application is addressed, can be launched by the user with only a small initial velocity and unlike a rifle the launching device need not be fabricated of high strength, relatively heavy, accurately machined elements. Instead, the miniature rocket of this invention can -be launched from any inexpensive and readily available device such as a soda straw, a cigarette, or any elongated tube suitable for imparting direction to the rocket. Because of the non-limiting character of the launching device, as will hereinafter be described, full and complete concealment of the rocket and launcher becomes possible thereby to render rockets more effective as a protective device, or as an offensive/defensive, or clandestine weapon of war.

Since the rockets are of such small dimension and weight, a series of rockets or a plurality of rockets can Ibe housed within a single launching device for sequential tiring, or for substantialy simultaneous firing, for coverage over troop concentration area. The entire assembly Weighs less than a fraction of a pound and occupies very little space.

Because the rocket gathers speed by acceleration in flight, as distinguished from acceleration within the confined length of a rie barrel, there is no resulting recoil or kick-back such that the rocket can be easily controlled and tired with very little, if any, previous training. It is only necessary for the operator to learn to aim the rocket weapon in the general direction of the target.

Since the rockets continue to accelerate during flight, the maximum velocity will be achieved in the probable target area at a distance from the point of launch. It becomes possible, with the miniature rockets of this invention, to develop velocities which exceed 3500 and even 5000 feet per second with speeds of better than 3000 feet per second obtainable 200 to 300 feet from point of launch. It is known that above certain projective velocities that the kinetic energy of the projectile is a more important wound ballistic parameter than momentum. Because of the extremely high velocities of these miniature rockets, obtainable in the probable target area, they are very effective kinetic energy kill mechanisms. Rockets have never before been used as a purely anti-personnel weapon, mainly because they have been too large for such a use; but also because they have not been accurate enough, they could not be made small enough, and they did not fly fast enough. They have therefore only been used to delivery payloads.

The rockets of the present invention are new and novel in that eminent scientists were very doubtful that the rockets could be made to fly reliably, nor that they could be made to ily fast enough for the intended use an anti- 3 personnel kinetic energy kill mechanism. These rockets produce new and unexpected results in that they have been made to fiy fast enough for the intended use, and theyare employed in an entirely new concept as a strictly anti-personnel lweapon.

These are but a few of the advantages capable of being derived by miniature self-propelled finned rockets which form the subject matter of this invention. Other advantages will become apparent hereinafter from the description of the invention.

This invention is thus limited to the important field of miniature rockets having a diameter within the range of -1/32 to l inch and a length within the range of 1/2 to 24 self-propelled rocketry as represented by rockets of considerably larger dimensions which are not designed for, and are impracticable and incapable of `being used as antipersonnel kinetic energy kill mechanisms.

Miniature rockets of the type described have not been produced merely by scaling down rockets of substantially large dimensions which are designed for non-equivalent use. There are a number of reasons advanced, by those skilled in the art, for the inability heretofore to produce small diameter finned rockets of high performance as antipersonnel kinetic energy kill mechanisms.

Rockets of small diameter were believed subject to very low efiiciencies and very low specific impulse. This may have been the result of the inability to achieve uniform radial burning of the propellant whereby imbalances would develop in the thrust to introduce error in the direction of flight. Such non-uniform burning would also cause exposure of some portions of the motor casing to higher temperatures than other portions whereby it was necessary to make use of casings of greater wall thickness and Weight to prevent case bursting. This results in inefficient low performance rockets due to the added weight.

The forces existing, by reason of the rates of acceleration, were believed by others to tend to collapse the propellant and cause malfunctions, and the erosive burning of the propellant and expected to cause additional Inalfunctions.

Since the burnings gases are at temperatures in excess 'of 4000 F., it was believed that the rocket elements, such as the casing, nozzle, and the like parts might suffer from relatively instantaneous thermal loads thereby causing shock disintegration of the rocket components before burnout of the propellant.

Since larger finned rockets traveling at speed of over 2000 feet per second suffer from aerodynamic heating, it was believed that a similar condition would occur in small rockets and that the amount of aerodynamic heating would be sufficient to introduce serious malfunctions due to the thin lightweight casing -which is necessary if the rockets are to fly at high velocity.

It was also believed that miniaturization would increase the inconsistencies of ignition and burning Which would cause still greater undesirable effects on small rockets thereby to render them inconsistent and unreliable in use.

Experiences with inhibitors used in propellants for rockets of large dimension indicated an undesirable tendency by the inhibitors to inactivate a large proportion of the surface of the propellant so treated. In miniaturized rockets it was thought that low performance and inefficient utilization of propellant would occur by leaving large amounts of unburned propellant in the rocket combustion chamber.

Further, the skilled in the art were of the opinion that small accurately dimensioned propellant grains, and other components, could not be produced for efficient use in miniature rockets.

It was believed that the small nozzles required for rockets of such small dimensions would be subject to a rate of ablation such as would cause loss of pressure and cessation of burning whereby the essential high acceleration rates could not be achieved or maintained. When the nozzle ablates severely, or erodes erratically,

vthe rocket is incapable of maintaining direction thereby creating a very erratic rocket with a resulting large course deviation.

It was also believed impractical to make use of a finned rocket since such thin'tins would be required as would be unable to withstand the high stresses incurred during fight at speeds in excess of Mach 2.

Reliable pyrotechnic fuses and igniters of the small dimensions and operational certainty required for small rockets of the dimension to which this invention is addressed were not available and were lbelieved to `be technically beyond the skill of the art.

Rocket cases of the small dimensions, stability, and tolerances contemplated by this invention were not available and were believed to be incapable of manufacture. The requirement is for uniform, balanced, thin walled casings capable of withstanding thermal load and shock, high aerodynamic heating, and high internal pressures and aerodynamic stresses, for high performance.

It is an object of this invention to produce and to provide a method for producing miniature rockets having a length less than 24 inches, and more specifically in the order of 35 mm., and a diameter less than l inch and more specifically in the order of 3 mm., which are capable of construction in a simple and efiicient manner to overcome the many difiiculties and problems confronting the manufacture and use thereof as previously described, and it is a related object to produce, and to provide for producing, miniature size rockets of the type described which are formed with an internal grain which is an insideoutside burner or burns uniformly radially from the inner surfaces radially outward to the casing for substantially complete combustion of the propellant, and which generates gaseous pressures at a rate for controlled acceleration of the rocket with little deviation in direction; and which can make use of a nozzle that ablates at a known and controllable rate for pressure control during burning of the propellant to achieve proper speed and direction in flight; which embodies sufficient strength to maintain the elements in the desired relationship in the assembly notwithstanding the large forces brought into play by the high rates of acceleration, the high temperatures and thermal shock to which the elements are exposed in use, and the high pressures and aerodynamic forces that exist during flight; which is capable of acceleration to velocities in excess of 5000 feet per second, and which is capable of impact speeds in excess of 2000 feet per second at ranges of 200 yards or more; which can be launched singly or in numbers from very simple and inexpensive launching devices; which can Ebe concealed in use thereby to constitute an offensive, defensive, clandestine weapon; which is capable of use in the atmosphere or under water without loss of efficiency or reliability; which can be launched in series or in clusters to achieve a shotgun effect for more effective coverage; which can be manufactured in a simple and inexpensive manner from a large variety of materials; which can be adapted for use with various types of payloads, and which has acceptable shelf life and can be packaged for shipment to distant locations in convenient and inexpensive packages for wider dissemination and application of use; and it is a still further object of this invention to produce and to provide new and improved means and devices for launching miniature rockets of the type described.

Those and other objects and advantages of this invention will hereinafter appear, and for purposes of illustration, but not of limitation, embodiments of the invention are shown in the accompanying drawings, in which FIG. l is a side view, partly in section, of a rocket embodying the present invention;

FIGS. 2 through 4 are aft-end views of a rocket, such as that shown in FIG. 1, illustrating methods of attaching ns to the rockets;

FIG. 5 is a side view, similar to FIG. 1, showing an alternate method of constructing the rocket;

FIG. 6 is a side view, similar to FIG. 1, showing a second alternate method of constructing the rocket;

FIG. 7 is a partial sectional view of an ignition device;

FIG. 8 is a side elevation of a miniature rocket having a heavy nose for stabilization as well as skewed fins for imparting a rotational movement to the rocket;

FIG. 9 is a side elevation of a rocket in place in a launcher tube, the rocket having canard fins for guiding it in the tube;

FIG. 10 is a side elevation of a rocket in place in a launcher tube, the rocket having elongated fins which guide it in the tube; and

FIG. 11 is a side view of a nose plug to be used in the FIG. 6 rocket.

In accordance with the present invention, a miniature rocket is provided Which is particularly effective against personnel as a kinetic energy kill mechanism. 'Ihe miniature rockets of the present invention are extremely small by comparison with other rockets heretofore produced. Typically, the rockets are dimensioned to have a length within the range of 8 to 35 mm. and a diameter of 11/2 to 3 mm. and weigh from 40 to 300 mg. Large numbers of rockets can be packaged in a launcher to form an assembly of low weight and small size. iBy firing a large number of such miniature rockets simultaneously from a launching tube, a shotgun effect can be achieved making it only necessary to point the launching tube in the general direction of the target. Since no particular skill is required, the weapon is particularly adaptable to use in guerrilla warfare and Iby unskilled military personel. The maximum velocity can exceed 5000 feet per second and can be `achieved within 50 .yards of the firing point. A

velocity of 3000 feet per second can be maintained sy far as 200 yards from the firing point. 'Ihus the rockets of the present invention comprise an effective, inexpensiveweapon of light weight and small dimensions which can be employed by unskilled personnel; Similarly, a large number ofsuch rockets can be'carried by a'bomb, mis-` sile, or shell for subsequent dispersionand'ring to cover a large area. l

Referring now to the drawings, the basic elementsof a rocket 11 (FIG. 1) generally comprise a tubular plastic or* metal casing 12 of extremely thin wall section 13 meas-v ured in the thousandths ofv an inch, which embodies a main tubular section `14 of small diameter, a `,nose 16 at the front end of the tubular rocket casing adapted to form. an ogive to extend as a continuation ofthe tubular section,

a plurality, suitably 3 or :4, of ns 17 having a thinl cross the body is 1%.," in'diameier. 1f desired, the nose of the rocket can be adapted to carry a payload such as arhigh explosive, or CBR (chemical, biological, or radiological) agent. Y, l Y,

With this type of rocket, velocities up to Mach 3. have been achieved at burnout. The rocket can penetrate sev.-v

eral layers of plasterboard, steel sheet, or wood or substantial thickness, and it is capable of bone fracture.

In the preferred practice of this invention, novelty exists not only in the rocket per se having dimensions of the type previously described, but also in the means for construction of the rocket elements including the casing, the inhibitor and case bonding agent, the propellant grain, the ignition means, the nozzle, the nose piece and ns, and their method of assembly to form a rocket: also in the means for packaging the rocket for storage, 'shipment, and use, and the means for launching'the rocket The tubing can be closed at the forward end by any one of the typical metallurgical forming processes of impacting, spinning7 swaging, or coining to form the nose.

A unique and novel concept of this invention resides in the fabrication of the rocket casing of thin metal foil or sheet stock. For example, a tubular rocket casing having an internal diameter of 1A; inch and a length of 11/2 inches and a wall thickness of 0.002 inch can be fabricated of metal foil, having a width of 11/2 inches and a thickness of 0.0003 inch, by winding the foil about a 1/s inch diameter core of mandrel to form six to seven layers which may be secured one to the other to provide a laminated structure. Such laminations may be joined into an integral casing by precoating the foil with a pressure-sensitive adhesive, a quick setting adhesive, or a heat setting adhesive, or by metal brazing, or gas welding, or by mechanical working for interfacial bonding. In the alternative, the multiple layers can be joined by electric or ultrasonic welding to produce a body having strength which exceeds the strength produced from a solid tube of the same metal. By the use of ultrasonic welding techniques, which are fast and relatively inexpensive, a high strength tubular casing, having an internal diameter of 1A; inch, and

a wall thickness of 0.002 inch, can be fabricated of 0.001

inch metal foil wound for two complete revolutions about a 1/8 inch diameter mandrel.

In the fabrication of the tubular casing by the described process of lamination of thin metallic foils, it is possible to incorporate thermal barriers in the casing between the foil layers, which, at the same time, increase the strength of the casing andthe interbonding of the lamina. Such modifications may be achieved by interleaving with thin non-metallic films of low heat conductivity. Such construction not only increases the strength of the casing but it minimizes the effects of thermal shock and enables the rocket to run cool for its short burn time thereby enhancing its resistance to bursting from the internally generated high temperatures and pressures.

Tubular-v casings, with or without fins, can also be formed by conventional techniques into tubular members of thicker wall section whereafter the wall sections can be scaled down to the desired dimension by conventional mechanical processes such as milling, chemical milling processes, or by electro-forming with the metal casing constituting one of the electrodes from which metal is taken into the electrolyte. v

v A tubular casing of the desired dimension and mass can also be fabricated by electro-forming techniques, using such metals as nickel, to produce a casing of exceptionally high strength.

A tubular casing of the desired dimension and mass can also be produced by impact extrusion. Use is made of aV small billet of the desired metal and it is impacted with a male member while at ambient low or high temperature to effect extrusion of a cavity of the desired dimension followed byheat treatment and/or coining, if desired.

When formed of a plastic or resinous material, with orrwithout fibrous or ller material, the tubular casing can be fabricated, with or without fins, by injection, impact, pressure, or by transfer molding techniques. Tubular metal combustion chambers can be inserted inside the plastic body to contain the pressure rise of fast burning The tubular rocket casings of the described dimensions can be fabricated of such materials as carbon steels, stainless steels, titanium, magnesium, aluminum and alloys of aluminum, brass, nickel, uranium alloys, plastics and filled, or loaded, resinous or plastic materials.

The nose of the rocket can be formed in various ways and the choice depends somewhat on the particular nozzle and fin configuration and/or structure to be used. If a nozzle is lightweight, e.g., formed directly as part of the casing, or is a lightweight insert, then a heavy nose is not required for proper balance and the nose can be formed simply as part of the casing by deep drawing, rolling, etc. But, if a heavy nozzle is formed or locked into the aft end of a rocket case or heavier than normal fins are used, then in order to counter-balance the weight, a heavy nose is required. A nose can be formed as a separate portion and locked into the tubular casing by metal forming techniques or the nose can be integral to the casing and weight added through the nozzle end. In the case of plastic rockets the weighted nose can be a needle pointed metal insert.

The rockets of the present invention may be stabilized in flight in various ways, preferably by the use of ns. It is desirable to provide a rocket in which a plurality, suitably three or four, fins extend radially outward from the tail end portion of the casing to guide the rocket in flight. Three fins are easier to effect for mass production, and three fined rockets are more easily packed in clusters, but four finned rockets are more stable in fiight. The fins can have a constant 24 (FIG. 2) or tapered 26 (FIG. 3) cross section from the root 27 to tip 28. It is desirable to have as sharp a leading edge 19 as possible to lessen the effect of the pressure wave on the fins which results from the passage of the rocket through the medium. The leading edges of the fins can have a tapered leading edge (FIG. 2) or the fins can be canted or skewed (FIG. 8) to impart a turning movement of rotational spin to the rocket in flight for minimizing the effect of imbalance in weight or burning rate.

FIGS. 2-4 show various ways in which stabilizing fins can be attached to the rocket. The fins can be secured to the case (FIG. 2) by means of welding or by adhesive. Three or more are attached individually to the casing. The fins can be formed as an integral portion of the casing (FIG. 3). This can be effected by electro-forming, upsetting, or drawing. Another method of forming the fins is shown (FIG. 4) wherein the lin 31 is made of two thicknesses of metal foil, 32, 33 both of which form a continuation of the metal body 34. In one method of forming the fins, a truncated cone is first drawn and the base of the cone is then creased longitudinally to form the fins 31. This provides a fin leading edge of folded foil 36. Alternatively, the fins may be made from simply a strip of metal foil. The complete fin structure is stamped and folded in one operation from the foil strip, then abutting interior portions of the vertical fins 32, 33 are glued together at the interface 37 and the fiared base portion 38 is glued to the rocket case. These fins can then be trimmed off to any desired shape and provide a very rigid lightweight and effective fin structure having a sharp leading edge 39.

The fins can be fabricated to various shapes, such as full delta, clipped delta, square, etc. The height of the fins should be sufiicient to maintain the fins in the air stream at speeds corresponding to high Mach numbers whereby they are not rendered ineffective by aerodynamic perturbations caused by the nose of the rocket. The shape and number must be adequate to provide aerodynamic stability. The mass of the fins should be as low as possible since in a fin stabilized rocket, as it is also desirable to maintain the center of gravity as far forward in the rocket as possible. The thickness and rigidity of the fins affect the drag and accuracy of the trajectory while their location on the rocket determines the static balance of the projectile and the size of the fin which must be used. All of these factors must be selected to achieve the highest possible accuracy and velocity for the rocket. To this end considerable investigation has been conducted with respect to the construction of fins; particularly with respect to the materials of which they can be formed, the manner in which they may be affixed, and the parameters of size, shape, number, and location on the casing.

The fins can be fabricated of very thin metal foil, such as aluminum, nickel, steel, or the like, to provide fins having a thickness measurable in a few ten thousandths of an inch such as 0.001 to 0.0005 inch in steel and a couple thousandths for aluminum. Such thin metal fins are preferably secured in place on the rocket casing by ultrasonic welding. Alternatively, suitable attachment can be achieved by cementing the fins onto the casing with a heat or pressure sensitive adhesive, epoxy cements, or by low temperature brazing or localized welding.

Fins can also be molded or cut of plastic film or sheet material and cemented in place on the rocket casing, or extruded as radially extending varies on a plastic tube which can be cut to length, slipped on the rocket case, cemented in place, and trimmed to shape.

When the rocket is formed with fins of thin steel or aluminum, the assembled tins can be sprung for wrapping circumferentially about the periphery of the rocket casing whereby the rocket occupies considerably less space in storage, packaging, launching, and handling. When released, the ns will normally spring back into shape to extend radially from the casing for use as stabilizers in flight. In the event that a slight curvature is retained by the fins, these function beneficially to cause the rocket to roll in flight and thereby obviate imbalances.

In a rocket in which the tubular casing is formed by the lamination of thin metal foils, as previously described, the fins can be formed integral with the metal casing by formation of the fins from the metal foil making up the outer layer or layers of the casing. For example, in a rocket system wherein the fin height is equal to the diameter of the rocket casing, as in a one caliber fin, the three fins can be split from the outer layer of metal foil and then folded radially outward to extend substantially perpendicularly from the rocket body. When the fins have a height which is less than one caliber, more than three fins can be formed and the fins can be arranged to extend radially from the casing in equally spaced apart relation. In the preferred practice, the fins are formed only of the outer foil of metal thereby to provide for minimum thickness with firm connection to the casing by reason of its integral attachment thereto. The fins can be folded out after the casing has been formed or the fins can be struck from the outer foil during formation of the casing.

The tubular casing 13 (FIG. l) can be lined internally with inhibitor and a case bonding material (not shown) at the interface between the propellant grain and the casing. Generally, the layer need only be 1/2 to 1 mil thick. The rockets can be flown successfully without inhibitor if uniform longitudinal ignition of the propellant grain is achieved whereby the propellant grain burns uniformly and does not expose one portion of the interior of the rocket case to the heat of combustion or flame front before nearly all of the propellant is burned.

Achieving uniform ignition with consistant regularity is difficult and therefore a very thin layer of inhibitor is generally employed to keep the heat from reaching the case. It was quickly determined that the standard inhibitors used for large rockets are unsatisfactory for miniature rockets in that they penetrate into the propellant and render a portion or layer of it useless. This unburned propellant is dead weight in small rockets, lowering the rockets efficiency, in addition to taking up valuable combustion chamber space which is at a premium when working with rockets of such small size. The standard inhibitors which are believed to have the effect of killing part of the propellant are those which use solvents of acetone, ether, ketone, methyl ethyl ketone, or other highly volatile solvents. Some inhibitors which were tried for the rockets of the present invention have been epoxys and polyester resins. These Worked but had relatively poor bonding characteristics. Plain rubber cement was tried, but the solvent was objectionable due to its inflammability. Titanium dioxide in a latex soluble base paint worked relatively well and it is believed a magnesium oxide soluble base paint would also work. The best inhibitors have been cellulose acetate, in the form of a spray, dip, or tape, ethyl cellulose, and polyvinyl acetate, polyvinyl alcohol, or polyvinyl chloride. A very thin layer is all that is necessary and if it is in the form of a tape which will bond to the propellant it is perfectly acceptable providing it is thin enough. It is possible that many types of tapes having a pressure sensitive surface which will bind to a high explosive surface may prove adequate and a simple solution to the problems.

The propellant grain is generally of a center burning configuration and is positioned inside the casing adjacent the outer wall extending radially inward terminating short of the central axis thereby providing a central bore through which the combustion gases issue during outward radial burning of the propellant from the inner surfaces to the case substantially simultaneously throughout the length. End burning propellants have not proven very successful due to the localized heat concentrations on the case which causes burn through. The inside radial burners provide a layer of propellant which insulates the case from the flame front until the propellant is nearly expended. Inhibitor between the propellant and case also provides insulation but lowers the effective propellant mass to total mass ratio. Inside outside burning propellants have been tried and have been fairly successful because they burn so fast that the burn period is over and the pressure expended before the case is significantly effected by thermal shock and pressure load. These propellants are effective for short range rockets, but are not completely consistent due to casev burn through problems.

In the practice of this invention, it is preferred to make use of a propellant which is of the double base type and which is quite' similar in composition to powders identified commercially by such names as Bullseye, Red Dot, etc., with or without additives, although other homogeneous propellants can be used.

The propellant is preferably employed as a single monolithic tubular grain of relatively thick walled cylindrical configuration dimensioned to have an outer diameter corresponding to the inner diameter of the bore of the casing so as to be received in close fitting coaxial relationship therein.

. l The cylindrical section is formed into a tubular member to provide a central bore for burning the grain radially outward while the combustion gases are exhausted in their entirety through the bore for passage rearwardly from the rocket through the nozzle whereby the rocket is driven in flight at an accelerating speed.

The monolithic grain can be machined from a solid rod or it can be cast or extruded to the desired inside and outside diameters from a liquid or paste containing the powder particles in a suitable vehicle. It can be formed by hot pressing directly into the rocket casing whereby a preferred intimate mating relationship is made between the outer periphery of the propellant grain and the inner surfaces of the casing to minimize the necessity for having to utilize an inhibitor and adhesive at the interface. Micro extrusion techniques have been employed successfully in the preparation of propellants for high performance miniature rockets.

In the assembly of the grain into the tubular rocket casing, use can be made of a special technique wherein the rocket casing is heated to an elevated temperature in the order of 100 to 200 C. and the propellant grain is cooled to a lower temperature in the order of 0 to 80 C. at the time of insertion to take advantage of the high thermal coefficient of expansion of the grain whereby as the elements return to normal temperatures, the propellant grain becomes stressed in radial compression by the ernbracing rocket casing. This provides for a sealed relationship between the grain and the casing thereby preventing cracks, defects and passages in the grain for the flame front to proceed along to reach the case. This obviates some of the problems heretofore encountered in case burn through. In addition, the intimate compressive Contact between the grain and casing enables the grain to rely on the casing for support substantially uniformly throughout its surface thereby to minimize breakdown of the grain under the extremely high pressures and temperatures and acceleration forces existing in flight. Thus the grain is limited to progressive radial burning from the inside out for generation of uniform gaseous pressures in flight.

Alternatively, the monolithic grain can be bonded to the casing by the use of a suitable adhesive which will adequately tie in the grain to the casing for support and which will also serve as a retardant or inhibitor to prevent burning of the grain at the outer wall.

A nozzle 22 (FIG. l) is fixed in the rearward or aft end portion of the casing in abutting relationship with the rearward end of the propellant with the nozzle formed to curvilinear shape for pressure control of the combustion gases generated internally of the casing by the burning propellant and issuing rearwardly from the bore through the tapered nozzle.

Nozzles can be formed in the rearward end portion of the casing by the metallurgical processes of impacting, rolling, spinning, swaging, or coining to provide a rocket casing which includes one end adapted to receive a nose piece, the tubular casing forming the combustion chamber, and the integrally formed nozzle. If nozzles are made in this way then it is possible to have a front loading rocket. Rockets can then be manufactured less noses and stockpiled with an assortment of desired noses such as steel needle points, CBR carriers, or high explosives, etc., and the proper nose can be selected and inserted when the need is determined.

Unless the nozzle is integrally formed with and during manufacturing of the casing, as previously described, the nozzle can be separately formed of such materials as solid or powdered metals, plastics, ceramics, and/or synthetic resinous materials which are filled with such materials as asbestos, glass, metal, or other fibrous elements or fillers, depending upon the end use of the rocket, the rates of nozzle ablation desired, and the overall design specifications.

Separately formed nozzles can be secured in position in the tail end portion of the tubular casing by means of cements, brazing, welding, such as ultrasonic welding, or by mechanical forming, such as crimping, rolling, or swaging.

In order to nullify eccentricities or discontinuities in the rocket cases introduced during the manufacture, and to improve accuracy, rolls or twists 41 (FIG. 2) can be incorporated in the throat or the flared exit portion of the nozzle to impart rotational spin to the rocket while in flight. While fins are employed in the rockets represented by the preferred practice of the invention as guiding members aud/or as turning members for the rocket in flight, fins are not essential when the rocket is rolled by means of vanes in the aft portion of the nozzle section. Roll can also be induced in the rocket by placing small chips or burrs on the divergent surfaces of the nozzle. This may be achieved by a burring drill, which is similar to a standard metal drill, that is positioned into the divergent exit portion of the nozzle and turned through a short angle of rotation whereby the material is fluted upwardly at an angle with respect to the surface to simulate two or more vanes.

Rocket nozzles of the types described can also be formed very simply and inexpensively by the metallurgical process of punching flat sheets of metal, plastics, or other sheet material, to the desired outside diameter. The part or parts can be punched, with or without coining, to form a conical cavity as well as the port. Punching or piercing operations of the type described yield a divergent or expansion cone to give an economically formed and eflcient nozzle.

The rocket can be fuzed by means which are only a few thousandths of an inch in diameter and are located within the cavity of the propellant grain, preferably extending to the forward end portion thereof. They iginite the grain uniformly longitudinally to enable the grain to burn efficiently and progressively radially outward. The fuzing of the rocket is extremely more difcult in miniature fin stabilized rockets of the type represnted by the practice of this invention where it may be desirable for a nite amount of time to lapse between ignition of the fuze and burning of the propellant. Such delays in fusing are desirable when the rockets are clustered in bomblets, or other delivery vehicles, to enable the rockets to be dispersed from the carrier and aerodynamically stabilized before ignition. Delay times of several seconds or even minutes are sometimes desired for some miniature rocket weapon delivery and employment systems.

A typical fuze can be made from a copper wire with a pyrotechnic coating, generally of a low ash material such as can be fabricated of propellant wool, granules, flakes, or powder. See copending applications Ser. No. 92,963, Rocket Fuse, led Mar. 2, 1961, and Ser. No. 95,391, Ignition Method, filed Mar. 13, 1961. This type fuze has a time delay value which is controlled by the dimension, composition, and configuration of the fuze. The ignition of the rocket motor is effected by a booster which is located at the forward end of the grain and is the internal terminal of the fuze.

Use can also be made of ignition means and an ignition train fuze system (FIG. 7) which burns at a predetermined rate through the bore of the grain from the nozzle end with the burning being maintained suiciently cool or insulated from the propellant so as not to ignite the grain. If a time delay is not necessary, electrical ignition can be used by employing small diameter enamel coated wires having the ends bared, twisted together, and embedded in a vglob of booster. A layer of inhibitor 1/2 to l mil thick, which can be provided on the internal surface of the grain (inside the perforation), prevents the pyrotechnic fuse from prematurely igniting the propellant before the booster effects uniform ignition of the grain. Inhibitor is particularly necessary in the case of time delay fuzes to prevent premature ignition.

The booster is provided at the forward end of the grain as a part of the fuzing system. While being minute in size it is capable of quickly liberating a large amount of heat suicient to uniformly and longitudinally ignite the grain. A special booster is used and is very important and a novel portion of the invention. It consists of a mixture of 36% boron potassium nitrate and 64% thermite as approximately 53% of the formula plus about 40% boron chlorate, 2% boron (9D-92% pure), and 5% nitrocellulose binder. The booster can be formed of small granules, akes, or powder. A rocket can be lighted by a deagrating fuze alone if it is hot and fast enough, but it must be very fast in order to light the propellant relatively simultaneously along it length to prevent subsequent case burn through and/ or end burning from uneven ignition.

The booster must be secured at the forward end of the grain within the combustion chamber. This prevents gases from the fuze from prematurely ejecting the booster out of the rocket and retains the booster in the combustion chamber for a time long enough for the propellant grain to start burning properly. Sometimes the fuze and/or booster will eject or become ejected from the exhaust end of the rocket before full ignition has been achieved to cause the rocket to abort. Stich losses may occur as the result of high acceleration during launch and burning. In order to prevent this from occurring the booster can be secured in the forward end of the combustion chamber by hooking the fuze over the grain to wedge 'it between the booster and the grain. Any other form of secure attachment will also serve the purpose.

Inhibitor or insulation prevents the fuze running to the booster from prematurely igniting the propellant in a cone or end burning configuration. It is an important parameter that the rocket be ignited with pure uniform radial burning. With delay fuzes, the interior portion may be inhibited for only a part of the length of the interior surface. The booster generates a hot re which sweeps back through the perforation along the grain penetrating the inhibitor layer and causing uniform ignition. The booster is in globular form, not necessarily in contact with the propellant grain. It must be of enough quantity to ignite the propellant through the inhibitor but not enough to detonate the propellant and thereby blow the case.

In the embodiment shown in FIG. l, the propellant grain can be cast directly into the rocket casing such as by centrifugal casting. This provides a uniformly molded dense grain intimately bonded to the case. In many instances this is desirable particularly since the nozzle of the rocket may be made of separate pieces and individually inserted into the `body after casting the grain.

In the embodiment shown in FIG. 6, the rocket has a casing which may be of extremely thin metal. A tubular member of the drawn metal, having a constricted rear portion, is manufactured separately and serves as the container for the grain. This embodiment is particularly advantageous since a simple rolling operation can be used to produce the nozzle and the tubular grain can be slipped into place in the casing from the nose and fastened therein by a suitable adhesive.

The propellant grains, such as illustrated, may be cast in place by inserting a hollow needle into the chamber and slowly withdrawing the needle while extruding a plastic propellant. Grains having a putty or powder-like consistency can be placed in the grain and formed by ramming a mandrel down the center, thereby compacting the grain to the walls of the case. The propellants can be of any of the well known types, but preferably are of a double base homogeneous type, such as smokeless powder.

In another modification, the rocket casing can be assembled with the propellant grain in place by precoating the metal forming the inner surface of the casing with a layer of propellant having a thickness desired in the grain. Thereafter the rocket can be formed as in the laminated system by turning about a mandrel dimensioned to have an outer diameter corresponding to the bore through the grain thereby to form the casing with the grain integral therewith. Under such circumstances, the grain can be formed in intimate bonded relationship with the inner wall of the metal casing so as to avoid the problems of providing an adhesive to bond the grain to the case and prevent burn through which might otherwise arise in the event of a spaced relationship between the grain and the casing. Or, conversely, the mandrel can be spray coated with propellant and the foil, having a pressure sensitive bonding agent applied to a portion of both surfaces, wrapped therearound.

A further representative embodiment of the invention is a plastic rocket 41 (FI-G. 5) dimensioned to have a diameter of approximately 1/8 inch and a length approximately ll/z inches. The casing 42 can be fabricated in one piece of a plastic such as polyamide nylon by injection molding. See patent application S.N. 93,209, Case Casting Method, led Mar. 3, 1961, by Bert B. Gould. one of the co-inventors of the present application. Four radially extending full delta fins 43 are formed as an integral part of the case and are dimensioned to have a thickness on the order of six to nine thousandths of an inch. The nozzle 44 is plastic, such as polyamide nylon, which ablates from the hot exhaust gases such that the erosion increases the cross-sectional dimension of the nozzle openings and prevents pressure buildup inside the casing by an amount which would otherwise burst the rocket.

A needle pointed metal nose 46 is inserted in the front end of the rocket case. The metal insert is necessary for a number of reasons. First, it gives the rocket a properly located CG (center of gravity) by counterbalancing the weight of the nozzle and ns. Even thou-gh the rockets can be own without heavy metal nozzles, as plastic ones have proved necessary for some uses, the metal nose weight is still necessary to counteract the weight of the ns and the nozzle material. The metal nose weight (needle pointed) provides for deep penetration of the target. Then, as the case portion of the rocket impacts, it is stopped due to its large size to weight ratio. The nose weight separates due to its high specific gravity and continues on into the target and creates considerable tissue damage due to tumbling.

The propellant 47 is of the double base type identified commercially by the term 1-12 or ARP, which is inhibited (not shown) with latex-titanium dioxide and extruded to they desired dimension to be received in fitting relationship within the casing. The fusing 48 is of 2 to 3 mil copper wire coated with a pyrotechnic mixture. A globule 49 of hot booster is placed on the end thereof and a portion of the front end of the propellant grain contains an amount of hotter burning pyrotechnic mixture to co-function as a booster. I'he nozzle is secured in the rearward end portion of the casing in abutting relationship with the rearward end of the propellant grain by adhesive. The plastic nozzleis of a particular ablating type having a straight section throat with a diverging nozzle for simplicity in construction. It was formed by punching on an eyelet type machine Whereas other nozzles can be formed as previously described, such as using a crimped section of tube 22 (FIG. 1).

A still further representative embodiment (FIG. 6) is shown. A thin wall casing 51 of extruded and redrawn aluminum or stainless steel is formed from a piece of tubing cut to the proper length. The front portion of the casing is preformed with a partially closed end 52 having a reduced diameter opening therein. A formed noseA plug 53 (FIG. 11), having a stud 54 projecting therefrom, is placed in the casing from the rearend of the rocket tube with the stud projecting through the reduced diameter opening at the front of the casing. The stud is then swaged around or spun over the reduced diameter portion of the casing forming an overlapping lip 56 on the casing and an aerodynamically streamlined nose to the rocket. The rear end 57 of the nose plug forms a bulkhead to the combustion chamber 58 which is the interior portion of the rocket casing between the nozzle 59 and the forward bulkhead. In theA present embodiment a thin walled center or radial burning propellant 61 is inserted in the casing. Before the propellant is placed in the casing the exterior surface can be painted or sprayed With a thin layer of inhibitor and/or bonding agent (neither shown) of latex, and/r ceramic mixture, using a water soluble binder. This inhibitor is at the interface 62 between the propellant grain and the inner wall of the casing to secure the grain in the casing and is applied to the grain in a manner to prevent exterior surface burning without attacking and killing a portion of the propellant. This keeps the flame off the case and prevents case failure and outside burning. Inhibitors, or bonding agents, or other materials may be used which prevent the collapse or disintegration of the grain during acceleration. Inhibitor is not used in the case of fast burning grains and inside-outside radial burners.

The fusing 63 comprises 2 to 3 mil copper wire coated with a pyrotechnic mixture. The fuse is provided with a hot booster 64 at the forward end of the propellant grain.

A ash sensitive substance 66 can be placed in the nozzle of the combustion chamber to seal it and provide lan initiator for the uze.

In assembly, the propellant grain, with the fuse and booster inserted, is tipped at its forward end with a fast setting resin and inserted with close tting relation into the casing. Care must be taken to prevent formation of bubbles Or open channels in the grain and inhibitor through which the high temperature combustion gases or the diame front might gain access to the casing. This is particularly important with a casing formed of aluminum since the strength of hte aluminum casing falls o rapidly when heated only to a temperature above a few hundred degrees Fahrenheit. Inhibitor is used on the interior surface of the grain to prevent premature and none uniform ignition.

A nozzle formed in a screw machine of carbon steel from solid stock having uniform thickness walls, conforming to the nozzle flow shape, is positioned in the rear end of the rocket casing abutting the propellant grain and is restrained within the rear end of the rocket casing by means of a curled restraining lip 67.

The fins 68 are fabricated of aluminum foil of 0.001 to 0.002 inch in thickness. In assembly, the aluminum strip Lfoil is coated with a heat-setting adhesive and the strips are preformed to corrugated shape. Thel formed ns are then held onto the rocket casing by means of a heated die which operates to set the ins on the casing and to bond the two thicknesses of foil together to form the fin which is subsequently trimmed to the delta shape.

The forward edges have a short vertical leading portion to insure proper bonding of the two n sheets and which is short enough that it does not extend up into the pressure wave, preventing its being bent over. The ns, to be effective, should extend approximately 1A: the length of the rocket case, and the vertical height should be about 1.4 calibers.

The CG of the rocket must be at least 60% of the distance .from the rear end of the rocket and in this embodiment it is about 63% of the distance. If the nozzles can be rolled into the case, the sameCG can be maintained for stability and the heavy nose eliminated. This permits a higher propellant to mass ratio and thereby higher maximum velocities are attainable.

In the embodiments of the invention the grain, fuzing, and boosters are interchangeable orv electrical ignition can be used where a time delay is not necessary.

The rockets can be molded or formed with canard fins 69 (FIG. 9). The leading edges 71 can be beveled to roll the rocket in ight. The canard fins serve as guide means for the rocket in a launching tube and thereby eliminate the necessity for any packing material in the bore of the tube. Alternatively, the ns may be made to extend the; length of the tubular portion of the'casing (FIG. 10) for supporting and guiding the rocket in the launch tube.A

The miniature rockets of this invention may be used for various purposes. The rocket is basically adapted forI use an an antipersonnel kinetic energy kill mechanism in the light of the high velocities which are reached shortlyz after launch and which can be maintained at ranges ofV hundreds of yards. The head of the rocket can be provided with a high explosive charge to impart tremendous physiological shock to the victim. Only very small amounts are necessary to effect this. The payload can be formulated to include CBR agents, incendiary materials, smoke generating materials, chaff for radar detection and jamming, or the like.

A plurality of such rockets can be packed in a bombletV preferably while the tins are folded circumferentially about the casing to conserve space. The rockets can also be packaged into containers or bomblets equipped with a ash or heat-generating mechanism which will ignite a ash or heat sensitive portion of the rockets. The flash sensitive compound is usually applied to the nozzle discharge portion and has the fuze embedded therein. The

compound receives the re and ignites the fuze initiating the tire train which may have a delay element incorporated therein. This ash sensitive compound, a new and novel composition, is formed of a mixture of approximately 66% potassium chlorate and 34% lead thiocyanate as 7580% of the mix plus approximately 20% zirconium powder and about 5% nitrocellulose binder. The proportions can be varied slightly without elfecting the sensitiveness of the compound.

In the above described example, the rocket casing is dimensioned to have a thickness of 0.005 inch. Such rockets can readily be accelerated to speeds of over 3300 feet per second in flight. When traveling at a speed of 2300 to 2700 feet per second, the rocket can penetrate over twenty layers of 1/2 inch Celotex with few, if any, distortions to the rocket body. When traveling at a speed greater than 2700 feet per second, penetration of the Same thickness of Celotex cannot be effected as the rocket body becomes crimped and bent on impact and sometimes tumbles. This causes a larger wound and increases antipersonnel effectiveness.

The following description will be with reference to the construction of a miniature rocket of metal foil in accordance Iwith the practice of this invention. For this purpose, use is made of steel or stainless steel foil of about 0.0003 inch in thickness. In the production of a rocket casing designed to have a wall thickness of about 0.002 inch, six to seven layers of the metal foil are wrapped about a mandrel, the outer diameter of which is dimensioned to correspond to the inner diameter of the rocket casing. When fully wound about the mandrel, the layers can be secured by a heat sensitive adhesive thinly applied to the foil.

A sharply pointed metal nose can be inserted onto the front end of the casing and fastened by crimp forming. The nozzle can be formed by the conventional process of swaging or co-ining the rearward end of the formed casing. The tins can be formed in advance of the nozzle by splitting or cutting the outer foil layer and turning the split portions outwardly to extend radially from the casing.

Before forming the nozzle, a monolithic tubuler shaped grain of a double base type propellant, inhibited with latex or water soluble resin, is extruded to proper form, dried, and inserted into the casing with the microprotechnic fuze and booster positioned inside the grain. The fuze can be insulated with plastic spaghetti or the grain inhibited on the interior surfaces for part of its length. The fuze is bared at the forward end and wrapped about the booster to anchor the fuze and booster within the propellant grain cavity during the unstable conditions of ignition and high acceleration.

Since the steel foil of which the casing is formed is resilient in character, the tins can be collapsed circumferentially about the peripheral surface of the rocket casing so as to occupy less space in loading or packaging. Any curl that remains after launch will cause the rocket to roll during flight for less dispersion. A rocket of the type described can achieve a velocity at burnout exceeding 3500 feet per second.

When foil of greater thickness is employed to form the rocket casing, the front end of the casing can be processed by impacting or coining to increase the mass density and to form a very sharp pointed nose. In this modification, the yfins are preferably separately formed of steel and ultrasonically welded in place onto the outside of the casing at the rearward end portion thereof. In the assembly of the rocket, the casing is heated to elevated temperature of about 100 to 200 C. while the propellant grain is chilled to a temperature within the range of 0 to 80 C. and the propellant, dimensioned to be received in ditting relationship within the casing, is inserted in the casing ywithout any inhibitor or adhesive.

The disc-like nozzle can be punched from carbon steel, to a conical shape for divergent surfaces and higher nozzle efficiency, as previously described. It can be assembled in place on the aft-end of the casing in abutting relationship with the rear end of the propellant grain assembly by crimping. A fusing means similar to that previously described can be employed.

The present invention offers a large number of advantages over the prior art of small arms weapons, such as the riile and bullets, in addition to those noted earlier.

The jets may be tired in various fashions. One may use an igniter and a firing button. The igniter leads to a detiagrating fuse, which has a plurality of branches. Each of the deflagrating fuse branches leads into the central cavity of one of the rockets. When the button is depressed, the igniter is set off, in turn ring all of the rockets simultaneously. Other ignition devices can be used such as electrical bridge wires or heater wires.

Due to the absence of any kick-back with rockets, use can be made of simple and inexpensive tubular members as launchers, for example, soda straws, bean shooters, cigarette wrappers and the like, with the ability to aim the rocket for a relatively high degree of accuracy at short ranges. This too permits substantially full concealment of the weapon for use as a defensive or clandestine weapon.

Use can also be made of honeycomb sections for storage of a multiplicity of miniature rockets, for support of the rockets in separated relation in the honeycomb sections, or for sequential or simultaneous tiring. The honeycomb, formed of paper, metal foil, or the like, thus functions either as a carrier and/or a launcher in which the rockets are each separately housed in separated compartments within a minimum of space.

The cases of the rockets can be made in a number of ways. For instance, they can be deep drawn, electroformed or extruded. Further, they can be made in more than one piece. The tube portions can be made individually, and the nose portion separately formed and fastened thereto. Further, the fins may be either formed in place or may be made separately and attached to the case, or they may be extruded with the case. Although the cases are preferably made of a thin strong metal, such as stainless steel, they can be made of plastic or combinations of metals, plastics, and resins, e.g., a plastic case can be combined with a metal combustion chamber.

The rockets of the present invention may be stabilized in ight in various ways including the use of the ns heretofore described. One method of stabilization is to make the nose heavy as by filling it with metal. Another method is to use skewed or canted tins or vaned cornbustion chamber nozzles so that a rotational movement or spin will be imparted to` the rocket in flight. Of course, various combinations of the stabilizing means can be employed.

Various guide means may be used to hold the rockets concentric until they leave the launcher tube. Thus, the rocket in the launcher is provided with short forward canard tins as well as rear fins. The lunching tubes can be provided with packing material which serves the same purpose.

Although it is generally preferred to` use standard rocket propellants, other propellants can be used, including high explosives. Thus, compounds such as HMX, which would normally detonate, merely burn when employed in small diameter rockets. Thus, compositions normally thought of as detonators can be employed as propellants and have the advantages of high density, high specific impulse, and high burning rates.

As has been pointed out above, the rockets of the present invention are well adapted for use by unskilled troops. The pattern formed by the rockets on leaving the launcher is such that accurate aiming is not necessary. The instructions for use are so simple that they can be given merely by cartoons printed on the launcher. By employing a carrier for the rockets together With dispersal and firing means, they form excellent weapons for covering a target of large area, such as dispersed troops. In addition to defense againstpersonnelthey can be used as anti-aircraft or underwater weapons and may be used to carry small payloads of CBR agents or high explosives. Since the rockets emit light for about the first 50 yards of their trajectory, they act as tracers aiding in aiming. Additional tracer material can be added to trace over the full range. From the foregoing it is obvious that there has been provided a unique small arms weapon, which can be manufactured in large quantities at low cost, which can 'be used by unskilled personnel, and which is highly lethal. v

As has been previously pointed out, the rockets of the present invention are of very small size and are of a different order of magnitude than the rockets heretofore proposed. Conventional rocket formulas do not apply to their performance so that they do not represent a mere scale down of existing rockets. For instance, the chamber pressure and burning rate formulas of conventional size rockets do not directly apply to them. For instance, according to conventional rocket formulas, the chamber pressure far exceeds the bursting strength of the case.

It will be understood that changes can be made in the details of construction and methods of fabrication of the miniature rockets without departing from the spirit f the invention. The invention is not to be limited, therefore, except as defined in the following claims.

What is claimed is:

1. A miniature rocket for use as an anti-personnel kinetic energy kill mechanism comprising in combination,

(a) a tubular casing capable of withstanding the pressure rise and thermal shock load of a high performance propellant with nose and rear end portions defining a combustion chamber therebetween,

(b) a weighted nose providing a center of gravity for said rocket at least 60% of the length of said rocket from the rear end and forming a forward restraining portion to said combustion chamber,

(c) a propellant grain having a central bore extending continuously axially therethrough lining the interior of said tubular casing abutting said forward restraining portion capable of accelerating said rocket to velocities in excess of 2500 feet per second and adapted to burn uniformly radially outwardly from the surface of the propellant bore to said casing with the gases of combustion exhausting rearwardly through the bore,

(d) means integral to said rocket for preventing the flame front of the burning propellant from reaching the case along its length before substantially all 0f said propellant grain is consumed,

(e) fusing means for providing relatively instantaneous uniform longitudinal ignition of the interior surface of the propellant grain, said fusing means further including a ybooster and means for securing said booster in said propellant bore, said booster comprising a mixture of approximately 3,6% boron potassium nitrate and 64% thermite as approximately 53% of the formula plus approximately 40% boron chlorate, 2% boron (9C-92% pure), and 5% nitrocellulose binder,

(f) means being provided in said rocket for preventing end burning, core burning and premature ignition of said propellant grain,

(g) at least three tail fins securely axed to said casing and extending radially outward therefrom, and

(h) nozzle means having a constricted throat therethrough secured in the read end of said casing abutting said propellant grain and formed for constricting the exhaustion of the gases or combustion from said propellant through said port to raise the pressure within the casing after ignition of the propellant, said nozzle means forming a rear bulkhead for said combustion chamber and supporting said propellant grain for preventing the ame front of the burning propellant from proceeding between the grain at the forward restraining portion and at the rear bulkhead to the casing.

2. The miniature rocket of claim 1 wherein the means provided in said rocket for preventing end burning, cone burning and premature ignition of said propellant is a 1/2 to l mil non-propellant penetrating layer of inhibitor or at least a portion of the length of the surface of said bore.

3. The miniature rocket of claim 1 wherein (a) said tubular casing tapers into an ogive at the nose end having a concentration of material relatively compacted therein, and

(b) said nozzle means is integral to said tubular casing and comprises a reentrant extension of the tubular casing.

4. The miniature rocket of claim 1 wherein a needle pointed metal nose is inserted and Securely held in the nose end of said rocket case.

5. The miniature rocket of claim 1 wherein the nozzle is formed of a plastic material that ablates in response to the passage of combustion gases therethrough for controlled increase in the size of the nozzle opening to compensate for the increase in the volume of combustion gases generated during radial burning of the propellant.

6. The miniature rocket of claim 1, wherein:

(a) said casing is of a relatively thin wall construction having a tapered open end portion; and

(h) a metal plug inserted in said open end portion,

said plug having a length sufficient to extend rearwardly to a limited extent into said casing and including a nose extending forwardly through the open end portion of said casing, said forwardly extending nose capable of -being swaged over the tapered open end portion of the casing and thereby provide an ogive portion therefor.

7. A miniature rocket for use as an anti-personnel kinetic energy kill mechanism comprising in combination,

(a) a tubular casing capable of withstanding the pressure rise with thermal shock load of a high performance propellant with nose and rear end portions defining a combustion chamber therebetween,

(b) a Weighted nose providing a center of gravity for said rocket at least,60% of the length of said rocket from the rear end and forming a forward restraining portion to said combustion chamber,

(c) a propellant grain having a central bore extending continuously axially therethrough lining the interior of said tubular casing abutting said forward restraining portion capable of accelerating said rocket to velocities in excess of 2500 feet per second and adapted to -burn uniformly radially outwardly from the surface of the propellant bore to said casing with the gases of combustion exhausting rearwardly through the bore,

(d) a 1/2 to l mil non-propellant penetrating layer of inhibitor afiixed to the propellant grain at the interface between said casing and said grain, said inhibitor including a component selected from the class consisting of titanium dioxide, magnesium oxide, cellulose acetate, ethyl cellulose, polyvinyl acetate, polyvinyl alcohol, and polyvinyl chloride,

(e) a 1/2 to 1 mil non-propellant penetrating layer of inhibitor covering at least a portion of the length of the surface of said bore,

(f) fusing means including a booster for providing relatively instantaneous uniform longitudinal ignition of the interior surface of the propellant bore, said booster further including a mixture of approximately 36% boron potassium nitrate and 64% thermite aS approximately 53% of the formula plus approximately 40% boron chlorate, 2% boron (90-92% pure), and 5% nitrocellulose binder,

(g) means for securing said booster inside said pro- 19 pellant bore until complete uniform ignition of the propellant has been eiected, A

(h) at least three tail ns securely aixed to said casing and extending radially outward therefrom, and

(i) nozzle means constructed of an ablative material 5 having a constrcted throat therethrough secured in the rear end of said casing tightly abutting said propellant grain and adapted to constrict the exhaustion of the gases of combustion from said propellant through said port to raise the pressure Within the casing after ignition of the propellant, said nozzle means providing a rear bulkhead for said combustion chamber and supporting said propellant grain for preventing the flame of the burning propellant from proceeding between the grain at the forward restrain- 15 ing portion and at the rear bulkhead to the case.

References Cited UNITED STATES PATENTS VERLIN R. PENDERGRASS, Primary Examiner 

1. A MINIATURE ROCKET FOR USE AS AN ANTI-PERSONNEL KINETIC ENERGY KILL MECHANISM COMPRISING IN COMBINATION, (A) A TUBULAR CASING CAPABLE OF WITHSTANDING THE PRESSURE RISE AND THERMAL SHOCK LOAD OF A HIGH PERFORMANCE PROPELLANT WITH NOSE AND REAR END PORTIONS DEFINING A COMBUSTION CHAMBER THEREBETWEEN, (B) A WEIGHTED NOSE PROVIDING A CENTER OF GRAVITY FOR SAID ROCKET AT LEAST 60% OF THE LENGTH OF SAID ROCKET FROM THE REAR END AND FORMING A FORWARD RESTRAINING PORTION TO SAID COMBUSTION CHAMBER, (C) A PROPELLANT GRAIN HAVING A CENTRAL BORE EXTENDING CONTINUOUSLY AXIALLY THERETHROUGH LINING THE INTERIOR OF SAID TUBULAR CASING ABUTTING SAID FORWARD RESTRAINING PORTION CAPABLE OF ACCELERATING SAID ROCKET TO VELOCITIES IN EXCESS OF 2500 FEET PR SECOND AND ADAPTED TO BURN UNIFORMLY RADIALLY OUTWARDLY FROM THE SURFACE OF THE PROPELLANT BORE TO SAID CASING WITH THE GASES OF COMBUSTION EXHAUSTING REARWARDLY THROUGH THE BORE, (D) MEANS INTEGRAL TO SAID ROCKET FOR PREVENTING THE FLAME FRONT THE BURNING PROPELLANT FROM REACHING THE CASE ALONG ITS LENGTH BEFORE SUBSTANTIALLY ALL OF SAID PROPELLANT GRAIN IS CONSUMED, (E) FUSING MEANS FOR PROVIDING RELATIVELY INSTANTANEOUS UNIFORM LONGITUDINAL IGNITION OF THE INTERIOR SURFACE OF THE PROPELLANT GRAIN, SAID FUSING MEANS FURTHER INCLUDING A BOOSTER AND MEANS FOR SECURING SAID BOOSTER IN SAID PROPELLANT BORE, SAID BOOSTER COMPRISING A MIXTURE OF APPROXIMATELY 36% BORON POTASSIUM NITRATE AND 64% THERMITE AS APPROXIMATELY 53% OF THE FORMULA PLUS APPROXIMATELY 40% BORON CHLORATE, 2% BORON (90-92% PURE), AND 5% NITROCELLULOSE BINDER, (F) MEANS BEING PROVIDED IN SAID ROCKET FOR PREVENTING END BURNING, CORE BURNING AND PREMSTURE IGNITION OF SAID PROPELLANT GRAIN, (G) AT LEAST THREE TAIL FINS SECURELY AFFIXED TO SAID CASING AND EXTENDING RADIALLY OUTWARD THEREFROM, AND (H) NOZZLE MEANS HAVING A CONSTRICTED THROAT THERETHROUGH SECURED IN THE REAR END OF SAID CASING ABUTTING SAID PROPELLANT GRAIN AND FORMED FOR CONSTRICTING THE EXHAUSTION OF THE GASES OR COMBUSTION FROM SAID PROPELLANT THROUGH SAID PORT TO RAISE THE PRESSURE WITHIN THE CASING AFTER IGNITION OF THE PROPELLANT, SAID NOZZLE MEANS FORMING A REAR BULKHEAD FOR SAID COMBUSTION CHAMBER AND SUPPORTING SAID PROPELLANT GRAIN FOR PREVENTING THE FLAME FRONT OF THE BURNING PROPELLANT FROM PRECEEDING BETWEEN THE GRAIN AT THE FORWARD RESTRAINING PORTION AND ST THE REAR BULKHEAD TO THE CASING. 